FIG. 1 shows a ducted fan gas turbine engine 10 comprising, in axial flow series: an air intake 12, a propulsive fan 14 having a plurality of fan blades 16, an intermediate pressure compressor 18, a high-pressure compressor 20, a combustor 22, a high-pressure turbine 24, an intermediate pressure turbine 26, a low-pressure turbine 28 and a core exhaust nozzle 30. A nacelle 32 generally surrounds the engine 10 and defines the intake 12, a bypass duct 34 and a bypass exhaust nozzle 36.
Air entering the intake 12 is accelerated by the fan 14 to produce a bypass flow and a core flow. The bypass flow travels down the bypass duct 34 and exits the bypass exhaust nozzle 36 to provide the majority of the propulsive thrust produced by the engine 10. The core flow enters in axial flow series the intermediate pressure compressor 18, high pressure compressor 20 and the combustor 22, where fuel is added to the compressed air and the mixture burnt. The hot combustion products expand through and drive the high, intermediate and low-pressure turbines 24, 26, 28 before being exhausted through the nozzle 30 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 24, 26, 28 respectively drive the high and intermediate pressure compressors 20, 18 and the fan 14 by interconnecting shafts 38, 40, 42 which are coaxially and concentrically arranged along a principal axis 31 of rotation for the engine 10.
It is well known that the efficiency of a gas turbine engine can be generally improved by closely controlling the gap between the various rotor blade tips and the engine casing so as to minimise the leakage of air over the blade tips. To this end, seal segments are located radially outwards of the turbine blades and provide the boundary of the main gas path. The seal segments often include an abradable liner which provides an adaptable and close fitting seal with the blade tips. The abradable seals are adaptable in that they preferentially wear when contacted by the blade tips such that the separating gap is determined by the blade tip position experienced in use. This allows the gap to be controlled to a working minimum without fear of damage to the blade tips.
One type of known abradable liner comprises a honeycombed structure in which a network of honeycomb shaped cells is presented radially outwards of the rotor blade tip path for abrasion. Such abradable honeycomb liners (or lands) often include a sintered powder coating within the honeycombs which helps provide increased oxidisation protection and a better seal with the blade tip. The sintered material is also less dense than the alternative metal of the seal segment honeycombs. However, the sintered powder coatings make it more difficult to provide effective cooling to the liner surface which can lead to increased oxidisation and premature degradation and wear of such liners.
Cooling schemes for abradable liners are known. For example, U.S. Pat. No. 3,365,172 describes a turbine shroud cooling scheme which provides cooling air through small holes which are registered with the openings in the honeycomb liner so as to provide cooling air to the gas washed surface of the shroud. However, this method precludes the use of a sinter powder coating and can result in a cooling regime which does not suit the e of the component.
The present invention seeks to provide an improved cooling arrangement for an abradable liner.